FIG. 1 illustrates a twin spool turbofan aircraft engine. An incoming airstream 3 is initially compressed by a booster compressor 6 and is then ducted to a high pressure compressor 9 in which the air is further compressed and from which the compressed air is delivered to a combustor 12. In the combustor, fuel (not shown) is injected into the compressed air, ignition occurs, and the hot, high-energy gas stream 14 which is produced is ducted to a high-pressure turbine 15.
The impact of the high-energy gas stream 14 causes rotation of the high-pressure turbine (HPT) 15, which, in turn, rotates the high-pressure compressor 9 connected to it. The high energy gas stream 14 then impinges upon a low pressure turbine 18, causing it to also rotate, thereby rotating the booster compressor 6 and a ducted fan 21. The fan 21 produces a propulsive airstream 24 which provides most of the thrust produced by the engine, while the residual gas stream 27 exhausting from the low pressure turbine 18 provides supplementary thrust.
The high pressure compressor 9, together with the high pressure turbine 15, comprise one spool, which is commonly called the "core." The other spool includes the fan 21, the booster 6, and the low pressure turbine 18.
Region 30 of high-pressure turbine 15 is shown in more detail in FIGS. 2 and 3A. In those figures, a clearance 33 is shown between blades 36 of the high pressure turbine 15 in FIG. 1 and a casing (or "shroud" or "stator"), 39 which surrounds the turbine blades. It is desired to maintain this clearance 33 as small as possible, in order to minimize the leakage of gasses indicated by arrow 42. Such leaking gasses impart virtually no momentum to the turbine blades 36, and represent a loss in energy.
It may be thought that the leakage problem may be eliminated by the expedient of manufacturing the engine with a sufficiently small clearance 33 which limits leakage to an acceptable value. However, such is not the case, because several factors cause the clearance 33 to change during engine operation. Five of these factors will now be explained.
First, during acceleration of the high-pressure turbine 15 in FIG. 1 from a ground idle speed of approximately 6,200 rpm to a take-off speed of approximately 11,000 rpm, the turbine disc 45A and blades 36 in FIGS. 1 and 2 expand in diameter (dimension 48 in FIG. 1) because of centrifugal force. This expansion is commonly termed "elastic growth" and is illustrated by the drop in clearance in region 61 in FIG. 3. The centrifugal force is quite large, as an example will show.
Centrifugal acceleration equals w.sup.2 r, wherein w is angular velocity, in radians per second, and r is radius, in feet. 11,000 revolutions per minute correspond to about 175 revolutions per second. If the diameter 48 is two feet, then the radius, r, is one foot, and thus the centrifugal acceleration equals (175.times.2.times.pi).sup.2, or 1.21.times.10.sup.6 feet/second.sup.2. Dividing this value by the acceleration due to gravity, namely, 32.2 feet/second.sup.2, yields a centrifugal force of approximately 37,600 G's.
This large G field occurs immediately upon acceleration and causes the diameter of the turbine rotor to increase. The actual increase in diameter from ground idle speed to take-off speed can be 0.028 inches. Therefore, if the diameter of the shroud 39 in FIG. 2 stays constant, the rotor elastic growth tends to diminish the clearance 33, and there exists a risk that the blades 36 may contact the shroud 39.
The second factor is the increased diameter of the shroud 39 which occurs because of the increased pressure of the gas stream 14. The pressure increase occurs at about the same time as the acceleration of the rotor 45 occurs. A typical pressure increase is from 41 psia at point 40 in FIG. 1 during ground idle to 380 psia at take-off speed. This increase in pressure can cause an increase in shroud diameter (which is twice the length of radius 41A in FIG. 2) of 0.004 inches. The increase in shroud diameter is represented approximately by region 43 in FIG. 3.
The third factor is the thermal expansion of the turbine blades 36 in FIG. 2: the temperature increase of the gas stream 14 causes the blades 36 in FIG. 2 to increase in length 51. A typical temperature increase of the gas stream 14 from ground idle to take-off speeds can be from 1300 degrees F. to 2500 degrees F. This increase in temperature causes the length 51 in FIG. 2 of the turbine blades 36 to increase, and by as much as 0.025 inches. This thermal blade growth is represented approximately by the reduction in clearance in region 38 in FIG. 3. This increase in length further tends to reduce the clearance 33 in FIG. 2.
The fourth factor is thermal growth of the shroud, caused by the increased temperature of the gas stream 14, and which increases shroud diameter. However, the increase in shroud diameter is much slower than the three changes in dimension which result from the three factors discussed above, and is represented by the gradual increase in shroud diameter indicated in region 44 in FIG. 3.
The fifth factor involves thermal growth of the turbine disc 45A of the turbine rotor 45, shown in FIGS. 1 and 35. While the disc 45A is not subject to the hot combustor exhaust 14 in FIG. 1, it is however in the presence of hot air which has been bled from the compressor 9 of the engine.
Compressor bleeds are used in order to accomplish such tasks as purging the internal region 54 of the engine of lubricant vapors and other gasses. The compressor bleeds are at higher-than-ambient temperature, causing the turbine rotor 45 in FIG. 2 to gradually assume a higher-than-ambient temperature, and thus to expand. The expansion is gradual because the compressor bleeds are not so hot as the airstream 14 (the hottest compressor bleed available is approximately 1100.degree. F.) and because the thermal mass of the rotor delays rotor heating. The rotor thermal growth is represented by region 55 in FIG. 3.
Therefore, to repeat: the clearance 33 in FIG. 2 is affected by the following factors in the following approximate order. Initially, (1) rotor elastic growth occurs, followed by (2) casing pressure growth. Then, (3) blade thermal growth occurs, followed by (4) casing thermal growth occur. Subsequently, (5) rotor thermal growth occurs.
A specific example of these changes in dimension will now be explained with reference to FIG. 3. Engine acceleration begins at a time of zero seconds, as indicated. The clearance at start up is indicated by point 66, and is approximately 0.048 inches. After the elapse of about 10 seconds, a take-off speed of 11,000 rpm has been attained, as shown in box 68, and centrifugal growth of the rotor causes a growth to approximately point 71, thereby shrinking the clearance.
Within circle 90 a minimum clearance occurs which then increases as time progresses. Such a minimum is termed a "pinch point" and places a limit upon the minimum clearance 33 in FIG. 2 which can be manufactured into the engine. For example, if the rotor were designed so that its initial clearance were clearance 93 in FIG. 3, the rotor would follow dashed line 94 upon acceleration, and the rotor would strike the casing at point 96, which cannot be allowed. Clearances at conditions other than the pinch point are more open than required. Therefore, to reduce this needlessly large clearance, active clearance control is used to control the diameter of the casing 39 by blowing cold air onto the casing. As shown in FIG. 1A, fan discharge air 97A is ducted to the turbine casing, as indicated by arrows 97, and a valve 134 controls the amount of air blown onto the casing.
Several forms of the invention provide improvements to the types of active clearance control just described, as well as to other types.